Automatic braking system for aircraft



5 Sheets-Sheet 1 c. E. GALLAGHER AUTOMATIC BRAKING SYSTEM. FOR AIRCRAFTS July 13, 1948.

Filed Oct. 29, 1945 In van for Char/es E. Gallagher Attorney July 13, 1948. c. E. GALLAGHER AUTOMATIC BRAKING SYSTEM FOR AIRGRAFTS 3 Sheets-Shed 2 Filed 001;. 29, 1945 July 13, 1948.

C. E. GALLAGHER AUTOMATIC BRAKING SYSTEM FOR AIRCRAETS s Sheets-Sheet 5 Filed Oct. 29, 1945 Inventor Char/es E Gallagher A Horne Patented 'July 13, 1943 UNITED T QFF'I'ZCE AUTOMATIC BRAKING SYST-EM FOR AIRCRAFT This invention relates-invgeneral to brake sys- ,tems' and inparticular toa novel differential brake system for aircraft which functions autostop.

The arrangement-is such that when the foreand-aft axis of- -the aircraft is =aligned onla selected heading, whichheading would normally be-the longitudinal'axis of the runway, an equal amount of-braking force is applied. to each wheel.

Flowever should the heading For r the aircraft departlfrom-l the runway axis, the-braking force on one wheel wilLbe automatically rincreased While thaton the other wheel pairedtherewith'will be automaticallynand simultaneously decreased to the end that =the.=increased braking effect infthe one wheel will then bring 't-he-aircra'fthealding back into alignment with the runway. aids.

: It is present practice to provide aircratt and especially those aircraft having a high landin speedacharacteristic, with-brakes for the left and right frontlanding wheels. The brakes" foreach of these wheels arearranged' for operation iridependently by the-pilot, usually by meansof-auxil-iary pedals located. on. the main foot pedals by whichthe pilot controlsoperation of the aircraft rudder. After thepilot has brought the aircraft intoa landing on the runway strip and-the aircraft wheels have made contact. with the runway, the pilotthen-endeavors to -apply .an equal braking effect toreach of-the wheelshsothat the aircraft can be brought-tea stop .withoutveerin off the runway. I-Iowever, should the aircraft tend .to i turn either one way .c vo1- the other after i the brakes have been :applied, the pilot willincrease the braking effect on one offthe Wheels and simultaneously decrease the braking effect on N the other wheel inrorderto bring the aircraftback on the selected landin course. Thusjfor ex- --ample, vshould the. aircraft tend to turn to the landing. course;

It-is thus seen that by using a. manually applied difierential type v.o-f brakingdescribed above, the pilot is, .to .a certain extent, able to control the coursevof the aircraft af-terltlhe landing, wheels have touched .the ground, and fthe speed. of theaircraft. hasbeen reduced to such an :matically to maintain the alrcraft on a .preselected heading while it is being braked-to a F2 extent; that the latter cannot be controlled very wiell by the rudder.

Whil'eamanually QperateddifEerentiaI braking system may beisati's'factory for some aircraft -it is obvious that afully automatic differential braking 'syst'e'mposses'sesmah'y advantages Over one that is manually operated.

The. trend in aircraft design is towards larger and, of course, much heavier and higher: speed planes; ,As therweight and speed factors increase, thelanding speeds also increase. Thus braking ofthe plane--becomes more -diflicult to control .manually, since:- only a very small amount of rincorrectly applied brakingforce will-send the'pla'ne veering off the runway.

The invention is alsouseful vfor providing-automatic steering of the plane on landing and :prior to take-off.

Al-s'o in operations 'where it is' 'desired that a -p-ilotless plane be controlled remotely, 'suchiias byradioga braking "system'of the automatic type made in accordance with principles 1 of -this invention now makes it practical toland' the heavier type-planes Withhigh landing'speeds,

'Automatic differentialbrakmg systems -willalso be very :useful in blind landing operations where :only 'the landing course is known. The pilot'-can set the braking controls to the selected'landlng course andi the plane will then be automatically 'brak'ed to a 'stop'without veering or hunting from such course.

Then too, since the pilot of a present day-large aircraft i already I has so many-other'du'ties to -perform, the desirability of relieiringhimfrom the further dutyof applying the planes brakes manually becomes obvious.

The generalobject of'myinventionis therefore to provide an automatically operated braketsystern for aircraft so -that the plane may bebrought safely toa 'stop withoutveering or' oscillating from'lthe selected landing heading.

5A more specific object is to provide an 'automatically cperatdflbrake system in which the amount of braking-force appliedto each ofpaired landing wheels of *the aircraft is' controlled by themagnitude and senseof apick-oif'signalp ro- 'duc'edbyadir'ectional gyroscope mountedonthe aircraft. 1 "These-and otherobjects' 'of'the invention'will become more apparent from the" following detailed" description and the accompanying" draw- "ings .which illustrate preferred ex'nbodiments-v of the invention.

In the drawings,

Fig. 1 is a diagrammatic view of one embodiment of the invention;

Fig. 2 is also a diagrammatic view illustrating a modified construction of my novel automatic differential braking system; Fig. 3 is a circuitdiagram of the servo amplifler component which is used in both the Fig. 1 and Fig. 2 arrangements; Fig. 4 is an enlarged longitudinal section view of the transfer valve component of the system in Fig. 1; and

Fig. 5 is a longitudinal section, somewhat enlarged, of one of the secondary valves utilized in sure differential which is then applied to the left and right paired wheels 14 and I5, respectively. of the aircraft. Other than the wheels and brakes, no parts of the aircraft have been illustrated in order to simplify the drawings.

The directional gyroscope The directional gyroscope I0 is of conventional construction, the one illustrated being of the electrically driven type which includes a 3-phase motor-gyro element l6 connected via conductors l1, l8 and I9 to S-phase power supply mains consisting of conductors 2|, 22 and 23. A B-phase alternator 24 feeds the 3-phase mains.

In accordance with conventional directional gyroscope construction, the gyro rotor element i6 is carried by an inner gimbal ring with the spin axis of the rotor arranged in a horizontal plane. The inner ring is pivotally supported by an outer gimbal ring, the latter in turn being pivotally supported by the instrument housing. The gimbal suspension being conventional, no details of its construction have been thought necessary. The directional gyroscope I0 is also provided with any of the well known types of caging mechanisms which function to lock the inner and outer gimbals with respect to the instrument housing. The caging mechanism being conventional, only a knob 25 for actuating it has therefore been illustrated. This knob is customarily positioned on the front panel of the instrument housin The directional gyroscope instrument in is installed on the aircraft in such a manner that when the inner and outer gimbals are caged, the spin axis of the gyro rotor It is aligned with the fore-and-aft, i. e., longitudinal axis of the aircraft.

Gyroscope it! includes a pick-off device 26, also of conventional construction, for generating signals proportional to the amount by which the aircraft departs from the spin axis to which the gyro rotor element I6 is set for braking, as will be more fully explained hereinafter.

This device converts deviation from any selected landing course into an A. C. voltage, the magnitude of which is proportional to the amount of deviation, and the instant polarity of which is dependent upon the direction or sense of the deviation from the selected landing course.

The pick-off device 23 consists of two parts, a rotor 27 attached to the vertical or outer gimbal pivot and a stator 28 attached to the stationary instrument frame. The rotor 21 is composed of a group of dumb-bell shaped iron laminations 4 around the shank of which a coil 3| is wound. The stator 28 is composed of a ring of iron laminations around which two coils 32 and 33 are wound in opposite paths. The two stator coils 32, 33 are connected in series opposition.

The coil winding 31 on the rotor 21 is connected via conductors 34, 35 and potentiometer 36 to one phase (conductors 22, 23) of the 3-phase A. C. supply mains.

When winding 31 on rotor 27 is excited with alternating current from the A. C. mains 22, 23, it produces an alternating flux. When rotor 2! lies in the position shown in Fig. 1, this flux enters and leaves the middle of each of the stator windings 32, 33 so that the flux in one-half of each winding is equal in magnitude but opposite in direction to that in the other half. The flux being alternating, voltages will be generated in the stator windings 32, 33; but since there is an equal amount of flux flowing in opposite directions through each half of each winding, the voltage across each of the windings 32-, 33 will be zero because the voltage produced in each half of each Winding is equal and opposite. Thus, for the position of the rotor 21 shown in Fig. 1, there will be no voltage across terminals C--D, even though the rotor winding 3! is excited,

Now, should the rotor 21 be turned counterclockwise from the position shown in Fig. 1, its magnetic flux field, of course, rotates with it. Each of the two stator coils 32, 33 is then traversed by the same flux flowing in one direction throughout its entire length so that a voltage is generated across each of the windings 32, 33. These two windings being connected in series, the two voltages add across terminals CD, and terminal C becomes positive during the same half cycle that terminal A of the rotor winding 3| is positive.

Should the rotor 21 be turned 90 clockwise from the position shown in Fig. 1, the flux pattern linking the stator coils 32, 33 is such as to produce an effect which is opposite to that produced when the rotor 21 is rotated 90 counterclockwise. That is to say, terminal D, instead of terminal C, then becomes positive during the same half cycle that terminal A is positive.

Should the rotor 21 be turned only slightly away from the position of zero voltage output, it will be evident that more turns of each of the two stator coils 32, 33 will be linked by magnetic flux going one way than by magnetic flux going the other way, so that there is a net voltage across output terminals C-D, but not so much as there would have been had the rotor element 21 been turned 90. Thus, for example, should the rotor element 21 be rotated clockwise from the position shown in Fig. 1, through an angle of, say, 45, terminal D would have the same polarity as terminal A during the same half cycle, but the voltage across terminals CD would not be as great as when the rotor 21 is turned through a full 90.

The stator element 28 of pick-off unit 26 may be adjusted manually relative to the rotor 21 and serves a purpose which will be fully explained in another part of this specification. The structure for permitting such manual adjustment includes a sector gear 20 integral with stator 28 and meshed with a worm gear 29 that is turned by a crank 30.

The details for mounting the rotor and stator elements of the pick-off device 26 on the vertical gimbal pivot and instrument frame, respectively, of the directional gyro instrument I 0 have not been-- illustrated isince the arrangement: is well known tothe art, ;itbeing common atthe present time to use, a directional. gyro with this type; of-=pick offin automatic. pilotiapp I In addition to the directional gyro-instrument l.0..- and.--the. pick-off. device 25 associated-therewith ior producing a control s gnal in accordance with the magnitude andsense of the departure of the aircraft from the selectedlanding course, it has been found desirable to further include the rate. gyro instrument ll from which there -;is produced a control. signal that varies with the. retreat which the aircraft departs from the selected landing: course. Such a 7 signal increases the sensitivity of -thesystem,=forl when it is super-- imposed upon the. signal from the directional gyro-J0, any. tendency of the aircraft to oscillate orl fhunt about the selected landing course .is

greatly. reduced. h v p s The rate gyro II, which may be ofconventional. construction,.includes a gyro rotor 58 which may be. constituted by the rotor element o f..a- 3.-\phasemotor also operating from the 3- phase power. 'supplyfmains 2 I23, and :connected thereto by conductors 39, Gdandd'l. The rate gyro H is mounted in the aircraft so that. the spin axis of the gyrorotorlies in a horizontal planevathwartships, that is to say, normal to the fore-and -aft axis of theaircraft. The gyro rotor 38 turnsina housingwhich is supported to rotate about. the fore-and-aft axis of the aircraft. Hwever,. movement ofth-e gyro housing is'restrained by;adjustable centering springs. Any change-in the heading of the aircraft causes the rate gyro to precess-andto thereby tilt the housing inwhich it spins. :The housing will be tilted to the right for a left turn and to the ileft for' a turn-towards the right, the movement being close- 1y proportional. to: the aircrafts t rate of turn.

Thev shaft whichs pportsthe movable gyro housingalsoncarries the rotor element 52 of a second pick-oil, device43. The stator element Ml of the-second pick-off device 43 issecured to the outer and stationary "frame of the-rate yro instr-umentr Preferably, the second: pick-01f *device 43: is of thesame construction of the first pick-off device 26.- That is, the rotor 42 includes a coil 45 wound thereon which is connected via conductors 46,, ll and potentiometer 48 to conductors 22, 23 of the S-phase A. C. supply mains 2|-23. The stator 44- has two coils 52 wound thereon in. opposite directions and connected in series opposition, the same as. in pickofi device 26.

Thus, as the aircraft turnsofi theselected landing course, the rotor 42 of thesecond pickoff. device 43: rotates with the gyro housing and producesiat-theoutput terminals C-D. of the stator windings 5|, {:2 a voltage signal proportional to the. rate at which the aircraft is turnmg. w

The details of the construction of the rate gyro instrument I I have-not been illustrated since theyare well known in the. art, it being common at the present timeto incorporate a gyro which puts out anelectrical signal in accordance with the rate-of-turn offth'e aircraft in automatic pilotisystems. As with the'pick-ofi device 26 associated with'the directional'gyro instrument l0, theamplitude of the signal voltage appearing acrossthebutput terminals C---Dv of the stator windings 51 and 52 will be prop'ortional'to the "amount byywliiohthe rotor l'2 is'turned' relative to:- stator .44 the-instant polarity of thee signat voltage will be? determinedbythedire'ction in which-therotor';42 rotates from its. -no-signal, or zero "voltage? position.

The output signal voltage from:pickofi device 26 associated with :thedirectional gyro instrument l ll is connected in series with the output signal voltage-derived: from piek-oif device 43" associatedwith the :rate gyro instrument H and. the two series-connected signal voltages are fed via conductors 53, 54 into;- the inputof a servo amplifierv unit 5'5.

It shouldialsobefnoted thatthe instant polarity. of theirate signalf fromtrate. gyro H will. be-the same as that of the signal fromfthedirectional.

gyro [0 when the turn hasthe samesense as the displacement. I However, as soon. asthe. differential' braking efi'ect stops the turning .ofthe. air-i craft, and-the displacement :signal then begins to decrease as. the aircraft comesback on v the selected heading, the turn-rate signal thenibecomes reversed. That is,-its instantpolarity is now opposite from th'atof'the displacement signal and the two series-connected signals will subtract from each other. Also connected in series withthese two. signal voltages is the output signal "voltage from a. third. piclr-o'ff d'evice- 5.6 which is actuated. by the hydraulic "difierential pressure follow-up;device [2. A more detailed description of thefollow-up unit l2 and pickofi 56 is included hereinafter.

Servo amplifier 55 is ,of conventional COD-1 struction and produces two electrical output voltages .which appear across the. three conductors 5'L'58and 59, the middle conductor 58 serving as a common conductor for both output voltages. The. arrangement is such that when the algebraic sumiof the pick-off signals derived from the threepick-ofi devices '26;'.43 and" 55 is equal to zero, the voltage. appearing between conductors 51 and 580i the output of amplifier 55 is equal to that appearing between conductors 58 and 59. However, under operating conditions such that the algebraic sum of these three signal voltages isgreater than zero. the voltage appearing' across output conductors 5! and'58 of amplifierfi55 will be greater or less, as the case may be, than thevoltage' appearing across output conductors 58'an'd 59.. Referring now tothe Fig. 3, the servo amplifier unit'55 is seen to'consist of two stages. Thefirst stage is constituted by a double. triode tube 62 which maybe a type6SL7.. The second stage of the amplifieris comprised of tWo'beam power tubes 63 and 6-4 "of the 2515 type,

The powersupply for the's'ervo amplifier unit 55 includes a single-phase transformer 65. This transformer hasa single primary winding 65a connected to conductors 22, 23 of the 3-phase supply'mains, and a'pair of secondary'windings 65b, 65c Whichare connected, as clearly shown, to supply. necessary voltage for the two anode circuitsof the firststage double amplifier tubeliz.

Asecond' transformer 66 is included to supply the necessary voltage for the anode circuits; of three rectifier tubes 61;.68 and 69 which may be of the 6X5type. This type tube is a'double idiodewith bothiha'lves operating in parallel. "he clearly shown, the primary 66a of transformer 65 is. connected in delta acrossthe f-phase supply mains 2|, 2'2and 2 3. The. primary windings are indicated as ELF andJG-P Two'secondari'es are located on the core of'each primary, Primary E and secondaries er and e2 are on the same core and therefore have voltages which are in phase with each other. Phases F and G are similar to phase E, the secondaries of these two phases being indicated by the reference numerals f1, and f2 and g1, and g2, respectively. The secondaries are connected zigzag, so that each line-to-neutral circuit has two secondaries'of difierent phases in series. Output leads I2, I3 and are connected to the anodes 61a, 68a and 69a of the three rectifier tubes 61, 58 and 89, and a neutral lead I5 serves as thenegative return. The cathodes 51b, 88b and 59b of the three rectifiers 61-69 are connected together, and through a 100 ohm filter resistor I6 to lead 11 and filter capacitor I8. Ihis combination forms a 3-phase full wave rectifier which produces a smooth D. C. output voltage, and which voltage is applied via leads I5 and I9 to the anode circuits of the second stage amplifier tubes 83 and 84.

Resistors 82 and 83 are connected in series across leads I5, I9 and their common point is grounded. The voltage drop across resistor 88 and its parallel capacitor 84 provides a positive bias for the grids 83b, 84b of the second stage amplifier tubes 63, 64.

The first stage of amplifier 55 also includes a potentiometer 85 and resistor 86 connected in series across the inputleads 53, 54, and the slider arm 85a of this potentiometer is connected to both control grids 52a, 82b of the double triode 62 by way of grid-current-limiting resistors 81, 88. Both cathodes 62c, 82d of amplifier 62 are connected together; they are also connected to input lead 54 through paralleled bias'resistor 9i and smoothing capacitor 92. As to operation of the first stage of amplifier 55, it is seen that when thevoltage appearing across input leads 53, 54 of amplifier 55 is zero, both control grids 62a, 82b of the double triode 62 are at the same potential with respect to the cathodes 62c, 82d, and this voltage is the drop across the bias resistor 9!. 'The secondary transformer windings 85b and 85c are connected so that at any instant, the voltage across one is exactly the reverse of the other. Since a vacuum tube conducts only when its anode is positive with respect to the cathode, thetwo halves of the double trio-de 82 will conduct during alternate half cycles, and have zero current during the other half cycle. In flowing through resistors 93, 94, these currents produce corresponding voltage drops directly proportional to the current at every instant. Both currents flow in the cathode circuit through resistor 9| and capacitor 92. The filter smooths this out to a smooth D. C. voltage which is positive at the cathode end and negative at the grounded end to which the grids 82a, 621) are connected. Thus, the grids have a negative bias which is the drop across resistor 9| and capacitor 92.

As to the second stage of amplifier 55, it is seen that the grid voltage of the second stage amplifier tubes 63, E4 is made up of three components. Starting from the cathodes 63a, 64a, there is first a negative drop in stabilizing resistor 95, caused by the passage of the cathode current through this resistor on its way to the rectifiertransformer neutral. .Changes in power supply voltage tend to change the current, and hence the voltage drop across resistor 95, but the changed drop changes the grid voltages in the direction to make the current return towards its original value. Thus, resistor 95 tends tostabilize the current at a constant value over a considerable range'of power supply voltage.

The stabilizing resistor 95 is connected to the negative D. C. supply lead I5, which is displaced from ground by the voltage drop across resistor 83 and capacitor 84. This introduces a positive voltage in the circuit from cat'hode83a to grid 63b, and cathode 64a to grid 64b, which is the second component.

The third component is the voltage drop across resistor 93 or resistor 94 in the first stage circuit, smoothing out to a direct current by capacitors 98, 91, 98, 99, and resistors I02 anl I03. This voltage component is negative.

With this circuit arrangement, and under a condition of zero voltage input signal voltage to amplifier 55, the grid voltages of the second stage amplifiers 63, 84 will be equal; the two tubes: then have equal'anode currents. That is to say, the voltage drop appearing between output leads 51 and 58 is equal to that appearing between leads 58 and 59, which constitute the two outputs of the amplifier 55 that are used to control the operation of the transfer valve unit I 3, previously mentioned. The nature of this control will be explained in further detail hereinafter in connection with that portion of the specification covering operation of the transfer Valve in the braking system.

The transfer valve The transfer valve I3 is provided for the purpose of converting the electric pick-off signals put out by the direction and rate-of-turn gyros, l8 and l I, into hydraulic signals in order to apply braking force to each of the landing Wheels I4 and I5. Valve I3 includes a cylindrical valve body I84 having a pair of high-pressure outlet ports I85 and I86, drain or low pressure return ports I81 and I08, and a high fluid pressure inlet port III.

The landing wheels I4, I5 of the aircraft are provided wit-h any conventional type of hydraulic brake structure which may comprise arcuate brake shoes '2, H3, and which are actuated by a piston movable in opposite directions within a hydraulic brake cylinder II4. A conduit Il5 connects outlet port I85 of transfer valve I3 to brake cylinder II4 on the right wheel I5. Similarly, the brake cylinder II4 of the left landing wheel I4 connects with outlet port I86 via conduit H6.

Slidable transversely within the-transfer valve body I04 is a stem II! which carries three valve port operators or pistons H8, H9 and I28. Stem I I1 extends out of both ends of the valve bodyIM and a solenoid is associated with each end of this stem.

Solenoid I23 includes a coil winding I24, a plunger I25 which is secured to one end of the stem Ill, and a spring I28 for maintaining the plunger I25 in a definite position with respect to winding I24. Similarly, the other solenoid I2' I includes an energizing coil winding I28, a plunger I29 secured to the other end of stem Ill, and a spring I38. Solenoid winding I24 is connected to conductors 58 and 59 which,.as previously explained, constitute one output from servo amplifier 55; winding I28 of the other solenoid I 2? is connected to conductors 51 and 58, constituting the other output from amplifier 55.

When the two outputs of amplifier 55 are equal in magnitude (which is the condition that prevails when the algebraic sum of the pick-oil voltages serving as inputs to the amplifier is zero), the forces acting upon plungers I25 and I29 are equal in magnitude but opposite in sense,

eases? with the resultthat the-three valve port operators I I8,-.-.I 2l)-a re maintained in a neutral position, which is the" position shown in Fig.1, Referring now to Fig. 4, it is clearly seen that when in a neutralposition,- the valve port operatorlIH is centered withrespect to the high pressure fluid inlet-part1 H, and that the latter is uncovered slightly an equal amount on both sides of the port operator H5, Thus, an equal amount of hydraulic-force is applied through the valve body I04 and via conduits H and IIS to the brake cylinder 4 of each of the two wheels I4 and I5, with the result that each of the wheels I4, I5 receives an equal amount of braking force.

Uncler'conditions where the pick-off voltages are not zero, a differential h-ydraulicforce is applied to the left and right wheel brahesthe differentialin braking force being in such direction as to reduce the pick-off signal voltages to zero. For example, let it lee-assumed that a pick-off alterhating current signal voltage appears across input leads 53; 54' of amplifier 55, and that such voltage is in,=phase with the voltage between conductors 22;,23 eithe 3-phasesupply mains (thatis, lead .53;;-is-,positiveduring the same half cycle that conductor 2,3 is positive). Depending uponthe settingof potentiometer, arm 850,, part =01 all of this signal-voltage will.- appear between ground and" the grids 62a, 62b of the double triode 62,

causing the anode currents through load resistors 83,1 94to become unbalanced. This, in turn; makes the voltage on grid 63b of tube 63 in the second stage of the amplifier more negative and the voltage on grid 84!) of tube-64 less negative. The-efiect pf this change in grid voltages on th v beam power tubes 63,64 is to unbalance their respective anode currents so that the current through solenoid winding I28 is; decreased,

while thencurrent through solenoid winding I24 is simultaneously increased. The magnetic pulls .of' the two solenoids I23, I21 are thereby/unbalanced, causing the valve stem I I1 tomoveto the right. This increases the pressure on the-fluid passedfrom; the high-pressure inlet port I II to the outlet port HIE-through conduit H5 and to,-

the brake cylinder H4 in the right wheel I 5. Simultaneously therewith, the pressure on the fluid passed from inlet port III to outlet port I06, conduit II6, and to: brake cylinder H4 in the lett wheel .I4' is decreased,

If -the instant polarity of the-pick off signal voltag -is reversed, that is,.lead 54 is positive during-the same half vcycle that conductor "23 on the 3"-'phase mains is positive, a reverse effect will take place. The brake pressure appliedto the left wheel I4 will then be increased, "while that applied to right wheel I5 willbe-decreased. A valve I09 is placed in the'highrpressure inlet conduit IHl' in order to releasethe pressure from the purpose of controlling the amountof the di-fierentialjnbrake ,pressures applied to the left and .righta-ircraft wheels I4,and I5, respectively,

forte. given .picli-ofi signal, and torestore equal braking pressures toithe .wheels" lf4, I 5' when the ,gyr .controls .,have made. such ;.correction as'has ,been -,necessary .to. bring l the fore and-aft aids thebrakes' when the latter are not in use,'as' fon of theaircraft back into alignment selected landing course or heading.

Included in the follow-up device I2 are a pair of, Bourdon tubes I32 and I33. The inner end of, tube I321 iscQnnectedvia conduit I34 into conduitIIG so that tube I 32 therefore becomes responsive to the hydraulic pressure in conduit I I76; Similarly,v the inner end of. Bourdon tube I33is connected via conduit I35 to conduit I I5; thus renderingtube I33 responsive to the pressure'iof the fluid incenduit u 5i The opposite ends ofBourdon tubes I32fand I133" are connected respectively to opposite ends of a transverse and floating link I36, it being noted that while the end'of tube I32'terminates in arfixed pivotat' the left end of link" I36, the end'of tube I33engages a forked portion 131184; the right end ofl his link; One .end'of avertica-l linhmember I38 is pivotally connected at I39 to. the cross link ['36, and the pposite end ot link I 3,8. is pivotally secured at L40 to a sector gear [41, the latter being mountedfor'rota'tion about a ,fixe'd: pivot ['42. The teethiin'f sectbr gear I 4| are meshed with pinion gear I43to'yvhich a shaft I14 isconnBCted. Shaft {44 rotates" the rotor" element 145* of" thejpicli -off unit 56', previlo usly mentioned. This. pick-offisv of fthefs'arne construction as .the other two pick- 0ft" units 2.5 and 43",Which'have been previously described. That is, pick-"off 5'5 includesa winding 1146' on the rotor element I45, such: winding'being congne'ctedvia conductors I41, I48 and potentiometer I49 to conductors 22;.23 ,on:.the'3 phas e, supply mains 2I'23j. The stator element. I51 of" the pick-off unit '56 includeswindings' J52, I5 3'that are arranged'in series. opposition. and connected in serie's at terminals :Cf D' with the outputs from the stator windings ofjthe pick-oi! units ZJBT and' :43 a pr viously xp a nedf It is thus ,se n .1 that wit h int onne te w t t e B t e t e a nka e. a ra ment de cribed,

any differential pressure which appears across thetwo Bourdon tubes I32 and? 133' will effect motion of the differential linkagemembers 136, I .3 8 and the sector1gearI4I- to thereby rotate the rotor element 1.,415Lolf" the .p clk-offfdevicet 5%6 throughan angleproportional to such ,diiferential the instantaneous. polarity of the pick-off signal produced by the pick-off'unit 56'is alwaysYopposite from that produced by the pick-,oiffdevice I that ,is vcontrolledJo'y the. directionaligylfQ vIll.

-With-.-this arrangement, it. thereforejevident that the amount of difierence in braking. effect 'onthetwo wheels 1.4 and I5 will-not .exceedthat which-is called for by a given-magnitude,oflpickoff-signalfromthegyro controls.

Operation of'Fz 1 system Assuminga landing operation; the aircraftiis first briought into' the runway with .the for'e-a'ndaft axis oftheaircraft inaligm'nent with the landing course selecte'd. The gyros l0 and II having been started sometime. previously are now ru-nnihg' at normal speed; and the .amplifier'BS is energized. The gimbal system of gyro I lliscaged.

When it is desired to apply the brakes' to .wheels I 4* and-'I5, the gyro-I0 is uncagedi by actuating lmob 25 andthe hydraulic supply turnedlon by closing switch I 2 25 The resulting action is the application of :pr'es'sur'e to thebrake shoeszIIZ'; I;I3

of both wheels I4 and I5 equally, thus retarding the aircraft. 7

The brake pressure will remain equal so long as the aircraft remains on the exact heading or course held at the instant the brakes were first applied. Under these conditions, the pick-off signal output from the pick-offs in gyros I and II is zero, as is also the output from the pick-off device 56.

Now, should the aircraft depart from the selected landing course, the pressure will be increased on the correct brake and decreased on the opposite brake in response to the pick-off signals generated and acting through amplifier 55 and transfer valve I3. The differential in braking pressures will be proportional to the algebraic sum of the pick-off voltages put out from the pick-off units 26 and 43 in accordance with the relative rotation between the stator and rotor windings of these Luiits, when the aircraft departs from the selected landing course. The amount of the differential in braking pressures is controlled by the follow-up unit I2, the pick-off unit 56 of which puts out a signal voltage proportional to the differential pressure and of opposite instant polarity to the combined signal voltages of the pick-off units 26 and 43, thus placing a balanced voltage on the servo amplifier 55 and centralizing the pistons I I8I20 in transfer valve I3. Thus any time that the aircraft departs from the selected landing course, the trans-- fer valve will operate and effect the necessary correction by establishing the differential braking forces. Actually, the transfer valve when in operation will be found to be oscillating at a rather high rate and holding a differential braking pressure strictly proportional to the changing conditions in the gyro signals.

The gradient of the pressure differential applied to thebrakes of the left and right Wheels I4, I may be adjusted to any desired slope by changing the setting of the potentiometers 36, 48 and I49 that are connected in circuit with the windings on the rotor elements of the pick-off devices 26, 43 and 56 previously described. .The slope of the differential braking pressure will also of course vary with the rate at which the aircraft departs from the selected landing course. That is, the magnitude of the pick-oif signal which is obtained from pick-off device 43 controlled by the rate gyro unit I I also becomes a factor in establishing the slope of the brake pressure differential.

The instant polarity of the rate signal output from pick-01f device 43 with respect to themstant polarity of the output signal from pick-off device 26 will be reversed as soon as the latter or displacement signal begins to decrease. The rate signal will therefore be subtracted from the displacement signal and hence cause a reversed braking differential to be set up as the aircraft approaches the selected landing course or heading thus resulting in a near critically damped control.

If it is desired to change the heading of the aircraft, say by 10, the turn may be effected by rotating the stator element 28 of the pick-off device 26 through 10. This is done by turning crank 30 which'rotates worm 29 to thereby rotate sector 'gear 20 that is attached to stator 28, as previously explained. When stator element 28 is thus rotated relative to rotor element 21, a displacement signal is generated at the output terminals of pick-off device 26. This signal is in addition to that generated from gyro action and thus a difierential braking, independent of gyro action, is established which thereby causes'the aircraft to turn. When the aircraft has turned through 10, and assuming no signal is then bein generated by gyro action, the stator 28 will have been turned through 10 in the reverse direction from which it was rotated, and hence be once again in a position as shown in Fig. 1 which is the no-signal attitude. Under such conditions, the transfer valve is once-again restored to the neutral position and the differential in braking: pressures disappears.

Reference is now made to the modified form of my invention shown in Fig. 2. The embodiment in Fig. 2 differs from that in Fig. 1 in that instead of establishing the braking pressure diiferential at the transfer valve I3, a pair of secondary valves I6I, I62 are utilized. The follow-up device is also modified. These valves are actuated by a servomotor I63 which is in turn controlled by the transfer valve. The construction of the transfer valve shown in the Fig. 2 embodiment differs slightly from that shown in Fig. 1 and, for this reason, has been designated by reference numeral I3 instead of numeral I3 which designates this valve in Fig. 1. For the same reason, the component parts of valve I3 have been given prime reference numerals. However, the directional gyro, the rate gyro, the pick-ofi devices associated with these gyros, and the servo amplifier components in the Fig. 2 embodiment are of the same construction as those in the Fig. 1 system and have therefore been given the same reference numerals as those used in Fig. 1.

It will be seen that piston H9 in the cylinder I84 of transfer valve I3 slightly overlaps the high pressure, inlet port I I I when the valve stem I I1 is in the center, or neutral position. Under this condition, the piston I64 of servomotor I63 will remain centered in cylinder I65 for the reason that the fluid pressure applied to the opposite sides of this piston from transfer valve I3 via conduits I66 and I61 are equal.

Piston rod I68 is pivotally connected at I58 to a plate I10, the latter being mounted for rotation about pivot I1I. Associated with plate I10, on opposite sides of pivot HI, and equidistant from this pivot, are piston rods I12 and I13 of the secondary control valves I6I and IE2. These piston rods include rollers I14 at their ends which are engaged with the sides of plate I10. Further, as shown clearly in Fig. 2, the rotor I45 of the pick-off device 56 is arranged to be rotated through pivoted link members I15 and I16, with one end of link I16 being connected to rotor I45 and one end of link I15 being connected to plate I18.

Referring now to Fig. 5, it is seen that each of the secondary control valves I6I, I62 includes a casing I11, a high pressure inlet port I18, a low pressure return port I19 leading to a sump (not shown) and an outlet port IBI. Supported within casing I11 for longitudinal movement is a tubular valve sleeve I82. Part of the latter is reduced in diameter to provide a passageway I83, which together with a valve head I84 function to control the flow of fluid from the high pressure inlet I18 to the outlet I8I which leads to the wheel brakes.

Motion of valvesleeve I82 is controlled by a rod I85 formed with a ball I86 on one end that is adapted to enter a conical seat I81 formed at the other end of sleeve I82 when pushed to the right from the position shown in the drawing. Rods I12, I13 to which reference has already been made are fixed to a cupped sleeve I88 that the rightendtoftheglatter "rests .agajinst' a collar l'93"formedjon'rdd' I85. "Thlls Tasr'd"I|2 or I13 isimoved in the directionhfithe arrow in Fig. '5, a force is applied,against thegleft' 'endfof .sprin'g I92. "The reactionproduced by the; springissuch as 'to 'cause a force'to' be, exerted at its" right endagainst collar-I931, pushing rod 185 to .theright sothat theba'lll'86 enters'seatl'lli'l. "Continued movement of rod I12 or I13 then causes: valve sleeve 182 to alsormove tojthe right sothat'the valve. head I84 is removed'irom its seat; When this occurs,."fluid can then pass from'the high pressure inletgllffl, through passageway, I88 to the outletpor'tfn'l. v

'VVh'en valves I 6 I I 62Tare assemhlefdfand' rollers IT4 are in. contact with. plate I10, the operating -.pa'rts of the .valves occupy such aposition that :the valve head I84 is always removed from its seat- "Thus,.'.the valvesf 61,1 fiz-are open, at all times. 'Vifith-the plate Ilil'fatthe neutral position, ass-shown in the 'drawingfthe two rods I12, I13 .occupy..l'ike,positions .an'dlthe valves are equally open, but as theIplate.)'fIUfismovedZin one, direction-.orLthe other. one of'l'theirods', either "I12 or I13, willbepushedin further whilethe other will becorrespondlngly 'force'dout to thus increase the opening ineonelofithe valves and-simultaneously decrease. the opening .id'f the other.

'The'hydraulic system o'fFig. 2 for supplying braking .pressure to the brakes on .wheels I4, I .includes avalve [96 l in. the main llngh .1 pressure supply. conduit .I'9I fthatis controlledby solenoid I98! and. switch 1'99; When'solenoidl I 98. is energized-by,.-c1osing-switch I99, the valve. I961 is in open. position. When de-energized,-.this valve iislclosed. Fluid..undenpresslireiasupplied to the -.in1et ports.of;valves. I 6.! [.fi-2.bymeans..ofi conduits .2-IlII-,Y2 l1I andl20'2 which-connect as.;shown .with theimain conduit J 91;

l'Phe. outletfromithei.auxiliaryscontrol valve .1161

is iconnectedlby conduit203 to the brakes. on the =right .wheel Il5. Branching iromiconduiti2fl3lls "another. conduitdfl l which :servesnas a byepasslto :the .low-.pressure.side ioresump of the hydraulic system (not shown). LA valve 2'l15- .controlled by solenoid] U-Biand. switch IiI- is .insertedinthe iconduit'firlllll Theoutlet .from thaauxiliary control valve I62 is .conhected by -conduit -2'08 .toithe brakes on the 1eft ..wheel I I 4;- Branching from .conduit (-208 .is .another conduit 209 whichalsorleadsltodthe sump of the hydraulic system...- Valve-2.I-I-controlled by solenoid 2I2 which is energized when switch 2 I 3-is closed .issinserted-r ins-conduit 20 9.

Valves205 and'ZII are closed when the braking system lSgplllTiIWGCOPBFflfIJiUl'lF- They are opened when it is desired to release the brakes, thus relievingall brake pressure from thehrakecylinders I I4 by opening branch conduits 204 menus to the sump.

It has been found desirableto also include by- Jpassnonnections to the sump ifrom thetconduits iconnectingthe cylinder I 045 withthe cylinder [-65 idf"servomotor I631 Thus, conduit I 66 iseiprov-ided vw'ithta' byl-passcondiiitfli I 6T and'valveLZl'Il-I=.that isoperatedbysolenoidu'la' which is energized when switch M9 is closed. Similarly, conduit I61 is provided with a by-pass conduit 22I and valve .222 Operated by solenoid 223 that is enerpressure inthecylinderof"th'eservomotor I63 to the'sump.

I Operation of Fig.2 system As with the system showninFig. 1-, after the aircraft has been brought into alignment'with thes'elected landing course, gyro-l li is uncaged byturn'ing knoh 2 5and switch l'flfl is closed. 'This energizes solenoid I98 which then opens valve I'Q'fi'to admithy'draulic fluid under pressure to theinlet portI-I-I 'of transfer valve l3." Switches 221; 2-I3, 2I9'and 224 'are opened so 'as to close li y-passvalves 211."), 2'I L2H, and '222. 'When the aircraft is aligned with the selected landing course,ithe output signals from the-gyro pickoffs- 26 and 43 will'be zero as previously-explained in connection with the Fig. 1 system,-and hence piston H9 in the transfer valve l3 will remain in a neutralposition as shown in the drawing. Under this condition the fluid pressure on both sides of piston I64 in cylinder I55'of'servo "I63 will be equal so that pistonlIi l is'alsomaintained in a'center position as shown in 'thedrawing. Plate I10 will therefore be retained in the position shown with the result that valves IG-I and I62 will be opened to'the'same extent. the fluidpressure attheoutletsof valves I6I,"I62 will be equal, and hence equal-fluid pressures will be transmitted through conduits 203 and 208 to 'thereloy apply an equal braking force to the brakeson wheelsl' l and I51,

Asin the Fig. 1' system; solong as the aircraft remains on the-selectedlandingjcourse, the brakthe fluid pressureon'one side ofpiston I64 and simultaneously decreases the ,pressure on the other side thereof. Piston-rod Ifi 8is thus 'calis'ed to-move an'd'rotate plate I-IIlfabout its pivot I'Ii. As plate I'Ifl frotates, say, inthe direction of the arrow," rod I"I2 will move in a direction indicated by, an arrow. to therebyreduce' the fluid pressure at; the outlet port IBI ofvalve IBI. Simultaneously, rod ITS will-move in a direction indicated by an arrow to thereby increase the v,fiuid pressure at the outlet port I8I ofvalve I62; Thus the'pressure inbrake line 208 and hencethe'brakingiforceon'leftWheel I4 will be increased while "the pressure in line 203 and braking force onright wheel '15 'is' simultaneously decreased. The increasing braking on the left wheel accompanied by asimultaneous"decrease in braking of the right wheel'turnsj the aircraft toward the left, putting it back'onthe selectedcourse. ISh'ould theaircrafttend to veer .towards'the-left, anopposite .efiect would takerpla'ce; i. 'e.,, the braking on the right wheel I 5' would be increasedand that .onthe 'leit wheel M woul'dbe, simultaneously.decreased.

The invention may' 'be embodiedlin other spegcific "forms without I. departing [from the .spirit or essential characteristics thereof. The present embodiments are therefore to be considered in all respects as illustrative and not restrictive. the scope of the invention being indicated by the appended claims rather than by the foregoing description, and all changes which come within the meaning and range of equivalency of the claims are therefore intended to be em braced therein.

The invention described herein may be manufactured and used by or for the Government of the United States of America for governmental purposes without the payment of any royalties thereon or therefor.

What is claimed is:

1. An automatic braking system for aircraft comprising, a plurality of braking means for said aircraft, a direction responsive device on said aircraft, a pick-off operated thereby to produce a signal corresponding to deviation of said aircraft from a preselected heading, turn rate device, a pick-off operated thereby to produce a signal corresponding to the rate-of-turn of said aircraft from said heading, and means controlled by both of said signals to establish a differential in braking effect in said plurality of braking means.

2. An automatic braking system for aircraft comprising, a plurality of braking means for said aircraft, a directional gyro on said aircraft, a pick-off operated thereby to produce a first sig nal proportional to the deviation of said aircraft from a preselected heading, a turn-rate gyro on said aircraft, a second pick-off operated thereby toproduce a second signal'proportional to the rate-of-turn of saidaircraf-t from said heading, means combining said first and second signals, and means controlled by said combined signals to establish a difference in braking effect applied to each of said plurality of braking means.

3. An automatic braking system for aircraft comprising, a plurality of braking means for said aircraft, a directional gyro including a first pick off operated thereby to produce a first signal corresponding to deviation of said aircraft from a preselected heading, a turn-rate gyro including a second pick-off to produce a second signal cor-responding to the rate-of-turn of said aircraft from said heading, means controlled by said first and second signals to increase the braking effect on one of said braking means and to simultaneously decrease the braking effect on another of said braking means, and a follow-up device including a third pick-off actuated thereby for producing a third signal corresponding to the difference in braking effect applied to said plurality of braking means, said third signal being arranged to combine with said two other signals to limit the differential braking effect to a preselected degree for a of said other signals.

4. An automatic braking system for paired landing wheels of an aircraft comprising, separate brake means for each of said wheels, a directional gyro, a pick-off operatedthereby to produce a first signal corresponding to deviation of said aircraft from a preselected heading, a turnerate gyro, a second pick-off operated thereby to produce a second signal corresponding to the rate-of-turn of said aircraft from said head ing, means combining said first and second signals, and means controlled by said combined signals to establish a difference in the brake efiect applied by said brake means.

5. An automatic braking system for aircraft comprising, a plurality of fluid operated braking given magnitude means for said aircraft, a directional gyro, a first pick-off operated thereby to produce a first signal corresponding to deviation of said aircraft from a selected heading, a turn-rate gyro, a second pick-off operated thereby to produce a second signal corresponding to the rate-of-turn of said aircraft from said heading, means combining said first and second signals, a source of fluid pressure, conduit means connecting said source to said brakng means, and a transfer valve interposed in said conduitmeans for establishing the amount of'fluid pressure applied to each of said braking means, said transfer valve being arrangedto be controlled by said combined signals.

6. An automatic braking system for aircraft comprising, a plurality of fluid operatedbr-aking means for said aircraft, a directional gyro, a first pick-off operated thereby to produce a first signal corresponding to the deviation of said aircraft from a selected heading, a turn-rate gyro, a second pick-off operated thereby'to produce a second signalcorresponding to the rate-of-turn of said aircraft from said heading, a source of fluid pressure, conduit means connecting said source to said braking means, a transfer valve interposed in said conduit/means for establishing a differential in the fluid pressure applied to each of said braking means, a follow-up device responsive to said pressure differential, said follow-up device including a third pick-off actuated thereby to produce a third signalcorresponding to said differential pressure, and means combinsignals.

7. An automatic braking system for aircraft comprising, a pair of fluid operated braking means for said aircraft, a directional gyro, a first pickoff operated thereby to produce a first signal corresponding to deviation of said aircraft from a selected heading, a turn-rate gyro, a second pick-ofi operated thereby to produce a second signal corresponding to the rate-of-turn of said aircraft from said heading, a source of fluid pressure, a transfer valve connected to, said fluid pressure source for establishing differential fluid pressure outputs, a servomotor actuated in accordance with said outputs, a pair of control valves actuated by said servomotor for increasing the brake pressure applied to one of said braking means and simultaneously decreasing the pressure applied to the other of said braking means, a follow-up device including a third pick-off actuated by said servomotor to produce a third signal corresponding to the position of said servomotor, and means combining all said signals, said transfer valve being arranged for control by said combined signal.

CHARLES E. GALLAGHER.

REFERENCES CITED The following referencesare of record in the file of this patent:

UNITED STATES PATENTS Number I I 2,382,727 Kronenberg'er Aug. 14,1945 

